The present invention relates generally to gas turbine engine shroud assemblies, and more particularly, to shroud assemblies having an inner surface machined to minimize blade tip clearances during flight.
Gas turbine engines have a stator and one or more rotors rotatably mounted on the stator. Each rotor has blades arranged in circumferential rows around the rotor. Each blade extends outward from a root to a tip. The stator is formed from one or more tubular structures which house the rotor so the blades rotate within the stator. Minimizing clearances between the blade tips and interior surfaces of the stator improves engine efficiency.
The clearances between the blade tips and the interior surfaces change during engine operation due to blade tip deflections and deflections of the interior surfaces of the stator. The deflections of the blade tips result from mechanical strain primarily caused by centrifugal forces on the spinning rotor and thermal growth due to elevated flowpath gas temperatures. Likewise, the deflections of the interior surfaces of the stator are a function of mechanical strain and thermal growth. Consequently, the deflections of the rotor and stator may be adjusted by controlling the mechanical strain and thermal growth. In general, it is desirable to adjust the deflections so the clearances between the rotor blade tips and the interior surfaces of the stator are minimized, particularly during steady-state, in-flight engine operation.
Stator deflection is controlled primarily by directing cooling air to portions of the stator to reduce thermally induced deflections thereby reducing clearances between the blade tips and the interior surfaces of the stator. However, because the cooling air is introduced through pipes at discrete locations around the stator, it does not cool the stator uniformly and the stator does not maintain roundness when the cooling air is introduced. In order to compensate for this out-of-round condition, the inner surfaces of the stator are machined so they are substantially round during some preselected condition. In the past, the preselected condition at which the stator surfaces were round was either when the engine was stopped or when the engine was undergoing a ground test. However, it has been observed that machining the stator so its inner surfaces are substantially round during either of these conditions results in the inner surfaces being out of round during actual flight. Because the inner surfaces are out of round during flight, the clearances between the blade tips and the inner surfaces of the stator vary circumferentially around the engine and are locally larger than need be. As a result, engine efficiency is lower than it could be if the stator inner surfaces were round during flight.
Among the several features of the present invention may be noted the provision of a method of machining an inner surface of a shroud assembly extending generally circumferentially around a central axis of a gas turbine aircraft engine. The engine includes a disk mounted inside the shroud assembly for rotation about the central axis of the engine and a plurality of circumferentially spaced rotor blades extending generally radially outward from an outer diameter of the disk. Each of the blades extends from a root positioned adjacent the outer diameter of the disk to a tip positioned outboard from the root. The method comprises determining a pre-machined radial clearance between the tips of the rotor blades and the inner surface of the shroud assembly during flight of the aircraft engine at each of a plurality of circumferentially spaced locations around the shroud assembly. Further, the method includes machining the inner surface of the shroud assembly based on the pre-machined radial clearances to provide a generally uniform post-machined radial clearance during flight between the tips of the rotor blades and the inner surface of the shroud assembly at each of the circumferentially spaced locations around the shroud assembly.
In another aspect, the present invention is directed to a shroud assembly for use in a gas turbine engine. The assembly extends generally circumferentially around a central axis of the gas turbine aircraft engine and surrounds a plurality of blades rotatably mounted in the engine. Each of the blades extends outward to a tip. The assembly comprises an inner surface extending generally circumferentially around the engine and outside the tips of the blades when the shroud assembly is mounted in the engine. The inner surface has a radius which varies circumferentially around the central axis of the engine before flight but which is substantially uniform during flight to minimize operating clearances between the inner surface and the tips of the blades.
In still another aspect, the shroud assembly comprises an inner surface spaced from a central axis of the engine by a distance which varies circumferentially around the central axis of the engine when the engine is stopped. The inner surface has a first locally maximum distance when the engine is stopped located at about 135 degrees measured clockwise from a top of the assembly and from a position aft of the surface. The first locally maximum distance is about 0.010 inches larger than a minimum distance of the inner surface. The inner surface has a second locally maximum distance when the engine is stopped at about 315 degrees measured clockwise from the top and from the aft position. The second locally maximum distance is about 0.005 inches larger than the minimum distance of the inner surface.
In yet another aspect, the shroud assembly comprises an annular support having a center corresponding to the central axis of the engine and a plurality of shroud segments mounted in the support extending substantially continuously around the support to define an inner surface of the shroud assembly. The inner surface is machined by grinding the surface to a radius of about 14.400 inches about a first grinding center corresponding to the center of the support, grinding the surface to a radius of about 14.395 inches about a second grinding center offset about 0.015 inches from the first grinding center in a first direction extending about 135 degrees from a top of the assembly measured clockwise from an aft side of the support, and grinding the surface to a radius of about 14.390 inches about a third grinding center offset about 0.015 inches from the first grinding center in a second direction generally opposite to the first direction.
Other features of the present invention will be in part apparent and in part pointed out hereinafter.